XFOIL Version 6.94 Calculated polar for: GOE 187 (SCHTTE-LANZ 2U10) AIRF 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.2582 0.01079 0.00285 -0.0388 0.6998 0.0804 0.500 0.3132 0.01057 0.00256 -0.0383 0.6711 0.0944 1.000 0.3905 0.00806 0.00251 -0.0432 0.6400 1.0000 1.500 0.4431 0.00827 0.00243 -0.0422 0.6024 1.0000 2.000 0.4958 0.00852 0.00246 -0.0413 0.5570 1.0000 2.500 0.5481 0.00887 0.00253 -0.0404 0.5039 1.0000 3.000 0.5997 0.00938 0.00270 -0.0395 0.4519 1.0000 3.500 0.6519 0.00988 0.00299 -0.0389 0.4192 1.0000 4.000 0.7046 0.01039 0.00339 -0.0383 0.3948 1.0000 4.500 0.7572 0.01087 0.00377 -0.0377 0.3711 1.0000 5.000 0.8101 0.01131 0.00421 -0.0372 0.3541 1.0000 5.500 0.8630 0.01175 0.00467 -0.0367 0.3364 1.0000 6.000 0.9154 0.01219 0.00516 -0.0362 0.3159 1.0000 6.500 0.9677 0.01259 0.00564 -0.0357 0.2884 1.0000 7.500 1.0626 0.01467 0.00724 -0.0336 0.1457 1.0000 8.000 1.1042 0.01645 0.00879 -0.0320 0.1028 1.0000 8.500 1.1486 0.01775 0.01010 -0.0305 0.0880 1.0000 9.000 1.1895 0.01930 0.01172 -0.0286 0.0791 1.0000 9.500 1.2273 0.02102 0.01351 -0.0265 0.0702 1.0000 10.000 1.2614 0.02289 0.01545 -0.0240 0.0606 1.0000 10.500 1.2936 0.02475 0.01752 -0.0213 0.0510 1.0000 11.000 1.3238 0.02652 0.01950 -0.0183 0.0422 1.0000 11.500 1.3415 0.02868 0.02181 -0.0140 0.0359 1.0000 12.000 1.3515 0.03163 0.02490 -0.0100 0.0326 1.0000 12.500 1.3595 0.03539 0.02892 -0.0073 0.0296 1.0000 13.000 1.3638 0.04007 0.03384 -0.0060 0.0274 1.0000 13.500 1.3674 0.04522 0.03914 -0.0057 0.0258 1.0000 14.000 1.3647 0.05141 0.04561 -0.0054 0.0245 1.0000 14.500 1.3608 0.05802 0.05252 -0.0061 0.0234 1.0000 15.000 1.3576 0.06450 0.05912 -0.0067 0.0224 1.0000 15.500 1.3525 0.07137 0.06621 -0.0069 0.0220 1.0000 16.000 1.3280 0.08201 0.07733 -0.0111 0.0216 1.0000 16.500 1.3013 0.09377 0.08951 -0.0162 0.0214 1.0000 17.000 1.2680 0.10769 0.10383 -0.0233 0.0213 1.0000 17.500 1.2278 0.12396 0.12049 -0.0325 0.0214 1.0000 18.500 1.1175 0.16920 0.16637 -0.0599 0.0220 1.0000